1. Field of the Invention
This invention relates generally to external cooling of turbine casings and flanges and, more particularly, to external cooling of low pressure turbine frames.
2. Discussion of the Background Art
Gas turbine engines, such as the General Electric CFM56-5B and -7B with dual annular combustors (DAC) engines, have been designed to operate both efficiently and with low amounts of pollution emissions. As a consequence, such engines have historically operated at very high EGT (Exhaust Gas Temperature) levels under ground idle and low power conditions. The turbine rear frame (TRF) and various nacelle components may experience temperatures that are high enough such that the mechanical properties of the materials are reduced and no longer acceptable from a fatigue and/or from an ultimate strength standpoint. To reduce these high temperatures, the present invention incorporates an external cooling manifold assembly capable of supplying sufficient cooling to this region of the engine such that temperature operating limits will be met or exceeded at all points of the flight mission.
Low pressure turbine (LPT) active clearance control systems which use an external manifold system to impinge fan discharge air on the LPT case have long been used for the purposes of maintaining desirable tip clearances between rotating turbine blades and respective surrounding shrouds, see U.S. Pat. No. 4,019,320, entitled "External Gas Turbine Engine Cooling For Clearance Control" as an example. Also long known in the art is external cooling of the engine case for the purposes of thermal control. Such cooling heretofore has been by flowing fan air over the casing during part or the entire engine operation.
The problem is particularly more acute for the rear turbine frame which typically supports the bearing assembly which supports the low pressure rotor and which contains the lugs through which pins are disposed to mount the engine to an aircraft pylon.